JP4108336B2 - Method and apparatus for reducing turbine blade tip temperature - Google Patents

Abstract

A rotor blade (40) for a gas turbine engine including a tip region (60) that facilitates reducing operating temperatures of the rotor blade is described. The tip region includes a first tip wall (62) and a second tip wall (64) extending radially outward from a tip plate (54) of an airfoil (42). The tip walls extend from adjacent a leading edge (48) of the airfoil to connect at a trailing edge (50) of the airfoil. A notch (80) is defined between the first and second tip walls at the airfoil leading edge. At least a portion of the second tip wall is recessed to define a tip shelf (90). <IMAGE>

Description

[0001]
BACKGROUND OF THE INVENTION
This application relates generally to gas turbine engine rotor blades, and more particularly to methods and apparatus for reducing rotor blade tip temperature.
[0002]
BACKGROUND OF THE INVENTION
Gas turbine engine rotor blades generally include an airfoil having leading and trailing edges, a pressure side, and a suction side. The pressure side and the suction side are connected at the leading and trailing edges of the airfoil and extend in a radial span between the airfoil root and tip. In order to facilitate reducing combustion gas leakage between the airfoil tip and the stationary stator component, the airfoil includes a tip region extending radially outward from the airfoil tip.
[0003]
The airfoil tip region includes a first tip wall extending from the airfoil leading edge to the trailing edge, and a second tip wall extending from the airfoil leading edge and connecting to the first tip wall at the airfoil trailing edge. Including. The tip region prevents damage to the airfoil when the rotor blades rub against the stator components.
[0004]
During operation, the combustion gas impinging on the rotating rotor blades transfers heat into the blade airfoil and tip region. Over time, continuous operation at higher temperatures can cause thermal fatigue in the tip region of the airfoil. To facilitate lowering the operating temperature of the airfoil tip region, at least some known rotor blades are slotted into the tip wall so that cooler combustion gases can flow through the tip region. Is provided.
[0005]
In order to help minimize thermal fatigue on the rotor blade tips, at least some known rotor blades include a shelf adjacent to the tip region to help reduce the operating temperature of the tip region. The shelf is formed in the pressure side of the airfoil and disturbs the combustion gas flow as the rotor blades rotate, thereby forming a film layer of cooling air against the pressure side of the airfoil. Make it possible. The film layer isolates the blades from the hotter combustion gases.
[Patent Document 1]
Japanese Patent Laid-Open No. 2000-297603
SUMMARY OF THE INVENTION
In an exemplary embodiment, a rotor blade for a gas turbine engine includes a tip region that facilitates reducing the operating temperature of the rotor blade without sacrificing the aerodynamic efficiency of the turbine engine. The tip region includes a first tip wall and a second tip wall extending radially outward from the airfoil tip plate. The first tip wall extends from near the leading edge of the airfoil to the trailing edge of the airfoil. The second tip wall also extends from near the leading edge of the airfoil and connects to the first tip wall at the trailing edge of the airfoil to form an open top tip cavity. At least a portion of the second tip wall is recessed to form a tip shelf. A notch extends from the tip plate and is formed between the first and second tip walls at the airfoil leading edge. The notch is in fluid communication with the tip cavity.
[0007]
During operation, as the rotor blades rotate, the hotter combustion gases near the leading edge of each rotor blade move to the airfoil tip region. As the tip wall extends from the airfoil, a narrow gap is formed between the rotor blade and the stationary structural component, which helps to reduce combustion gas leakage therethrough. When rubbing occurs between the stationary structural component and the rotor blade, the tip wall contacts the component and the airfoil remains intact. As the rotor blades rotate, lower temperature combustion gases near the leading edge flow through the notches, creating a cooler gas temperature in the tip cavity. Combustion gases on the pressure side of the rotor blades also flow over the tip region shelf and mix with the cooling air. As a result, the notches and shelves help reduce the operating temperature of the rotor blades in the tip region, but do not consume extra cooling air and thus improve turbine efficiency.
[0008]
DETAILED DESCRIPTION OF THE INVENTION
FIG. 1 is a schematic diagram of a gas turbine engine 10 that includes a fan assembly 12, a high pressure compressor 14, and a combustor 16. The engine 10 also includes a high pressure turbine 18, a low pressure turbine 20, and a booster 22. The fan assembly 12 includes a row of fan blades 24 extending radially outward from the rotor disk 26. The engine 10 has an intake side 28 and an exhaust side 30.
[0009]
During operation, air flows through the fan assembly 12 and pressurized air is supplied to the high pressure compressor 14. The highly pressurized air is sent to the combustor 16. Airflow from combustor 16 (not shown in FIG. 1) drives turbines 18 and 20, and turbine 20 drives fan assembly 12.
[0010]
FIG. 2 is a partial perspective view of a rotor blade 40 that may be used in a gas turbine engine, such as gas turbine engine 10 (shown in FIG. 1). In one embodiment, the plurality of rotor blades 40 form a high pressure turbine rotor blade stage (not shown) of the gas turbine engine 10. Each rotor blade 40 includes a hollow airfoil 42 and an integral dovetail (not shown) that is used to attach the airfoil 42 to a rotor disk (not shown) in a known manner.
[0011]
The airfoil 42 includes a first side wall 44 and a second side wall 46. The first side wall 44 is convex and forms the suction side of the airfoil 42, and the second side wall 46 is concave and forms the pressure side of the airfoil 42. The side walls 44 and 46 are joined at the leading edge 48 of the airfoil 42 and an axially spaced trailing edge 50 located downstream from the leading edge 48.
[0012]
The first and second side walls 44 and 46 extend longitudinally or radially outward, respectively, from the blade root (not shown) located adjacent to the dovetail to the radius of the internal cooling chamber (not shown). It extends in a span to the tip plate 54 that forms the outer boundary in the direction. A cooling chamber is formed inside the airfoil 42 between the side walls 44 and 46. Internal cooling of the airfoil 42 is well known in the art. In one embodiment, the cooling chamber includes a tortuous flow path that is cooled with compressor extraction air. In another embodiment, sidewalls 44 and 46 include a plurality of film cooling holes (not shown) extending through the sidewalls to facilitate additional cooling of the cooling chamber. In yet another embodiment, the airfoil 42 includes a plurality of trailing edge holes (not shown) that are used to discharge cooling air from the cooling chamber.
[0013]
The tip region 60 of the airfoil 42, sometimes known as the squealer tip, includes a first tip wall 62 and a second tip wall 64 that are integrally formed with the airfoil 42. The first tip wall 62 extends from near the airfoil leading edge 48 to the airfoil trailing edge 50 along the airfoil first side wall 44. More specifically, the first tip wall 62 extends from the tip plate 54 to the outer edge 65 by a height 66. The height 66 of the first tip wall is substantially constant along the first tip wall 62.
[0014]
The second tip wall 64 extends from the vicinity of the airfoil leading edge 48 along the second side wall 46 and is connected to the first tip wall 62 at the airfoil trailing edge 50. More specifically, the second tip wall 64 is spaced laterally from the first tip wall 62 such that a top cavity 70 with an open top is formed by the tip walls 62 and 64 and the tip plate 54. Arranged. The second tip wall 64 also extends radially outward from the tip plate 54 to the outer edge 72 by a height 74. In the exemplary embodiment, second tip wall height 74 is equal to first tip wall height 66. Alternatively, the height 74 of the second tip wall is not equal to the height 66 of the first tip wall.
[0015]
A notch 80 is formed between the first tip wall 62 and the second tip wall 64 along the airfoil leading edge 48. More specifically, the notch 80 has a width 82 extending between the first tip wall 62 and the second tip wall 64, and a first portion 86 and a bottom 86 of the notch 80 formed by the tip plate 54, respectively. And a height 84 measured between the outer edges 65 and 72 of the second tip wall.
[0016]
In another embodiment, the notch 80 does not extend from the tip plate 54, but instead from the notch height 84 from the first and second tip wall outer edges 65 and 72 to the tip plate 54, respectively. It extends a small distance (not shown), so the notch bottom 86 is at a distance (not shown) from the tip plate 54. In yet another embodiment, the second tip wall 64 is not connected to the first tip wall 62 at the airfoil trailing edge 50, and an opening (not shown) at the airfoil trailing edge 50 has a first tip wall 62. And the second tip wall 64.
[0017]
The notch 80 is in fluid communication with the open-ended tip cavity 70 to allow cooler combustion gases to enter the cavity 70 for low temperature heating purposes. In one embodiment, the notch 80 also includes a guide wall (not shown in FIG. 2) that is used to flow the flow entering the open top tip cavity 70 toward the second tip wall 64. More specifically, the guide wall extends from the notch 80 toward the airfoil trailing edge 50.
[0018]
The second tip wall 64 is at least partially recessed from the airfoil second side wall 46. More specifically, the second tip wall 64 is recessed from the airfoil second side wall 46 toward the first tip wall 62, generally between the airfoil leading edge 48 and the trailing edge 50. A first tip shelf 90 that extends radially outward is formed. More specifically, the shelf 90 includes a front end edge 94 and a rear end edge 96. The front edge 94 and the rear edge 96 are each tapered and are flush with the second side wall 46. The shelf leading edge 94 is located downstream of the airfoil leading edge 48 by a distance 98, and the shelf trailing edge 96 is located upstream of the airfoil trailing edge 50 by a distance 100.
[0019]
The recessed second tip wall 64 and the shelf 90 form a substantially L-shaped trough 102 therebetween. In the exemplary embodiment, tip plate 54 is generally non-porous and only includes a plurality of holes 106 that extend through tip plate 54 in tip shelf 90. The holes 106 are axially spaced along the shelf 90 and provide fluid communication between the trough 102 and the airfoil internal cooling chamber. In one embodiment, tip region 60 and airfoil 42 are coated with a thermal barrier coating.
[0020]
During operation, the squealer tip walls 62 and 64 are placed in close proximity to a conventional fixed stator shroud (not shown) to form a narrow gap (not shown) therebetween to reduce combustion gas leakage therethrough. To promote. The tip walls 62 and 64 extend radially outward from the airfoil 42. Thus, if rubbing occurs between the rotor blade 40 and the stator shroud, only the tip walls 62 and 64 will contact the shroud and the airfoil 42 will remain intact.
[0021]
Because the combustion gas exhibits a parabolic shape flowing through the turbine flow path, the combustion gas near the leading edge 48 of the turbine blade tip region is at a lower temperature than the gas near the trailing edge 50 of the turbine blade tip region. As cooler combustion gases flow into the notches 80, the thermal load on the tip region 60 is reduced. More specifically, the combustion gas flowing into the notch 80 is at a higher pressure and a lower temperature than the gas leaking from the rotor blade positive pressure side 46 to the rotor blade negative pressure side 44 through the tip clearance. As a result, the notch 80 helps to reduce the operating temperature in the tip region 60.
[0022]
Further, as combustion gas flows past the airfoil first tip shelf 90, discontinuities in the airfoil pressure side 46 provided by the trough 102 cause combustion gas to flow from the airfoil second side wall 46. Since it peels, the fall of the heat transfer is accelerated | stimulated. In addition, the trough 102 provides an area where cooling air accumulates and forms a film for the sidewall 46. The first tip shelf hole 106 discharges cooling air from the airfoil internal cooling chamber to form a film cooling layer on the tip region 60. Due to the rotation of the blades, the combustion gases outside the rotor blade 40 at the leading edge 48 near the blade pitch line (not shown) travel along the second side wall 46 and the airfoil tip region 60 near the trailing edge 50. The leading edge tip operating temperature is lower than the trailing edge tip operating temperature. The first tip shelf 90 functions as a backward-facing step in the moving radial flow, and shields the film of cooling air accumulated on the side wall 46. As a result, the shelf 90 helps to improve the cooling efficiency of the film and lower the operating temperature of the side wall 46.
[0023]
FIG. 3 is a cross-sectional view of another embodiment of a rotor blade 120 that may be used in a gas turbine engine, such as gas turbine engine 10 (shown in FIG. 1). The rotor blade 120 is substantially similar to the rotor blade 40 shown in FIG. 2, and components in the rotor blade 120 that are identical to the components of the rotor blade 40 are indicated using the same reference numerals used in FIG. 3 is specified. Accordingly, the rotor blade 120 includes an airfoil 42 (shown in FIG. 2), sidewalls 44 and 46 (shown in FIG. 2) extending between the leading edge 48 and the trailing edge 50, respectively, and a notch 80. Further, the rotor blade 120 includes a second tip wall 64 and a first tip shelf 90. In addition, the rotor blade 120 includes a first tip wall 122. The notches 80 are formed between the first tip wall 122 and the second tip wall 64, respectively.
[0024]
The first tip wall 122 extends from the vicinity of the airfoil leading edge 48 along the first side wall 44 and is connected to the second tip wall 64 at the airfoil trailing edge 50. More specifically, the first tip wall 122 is spaced laterally from the second tip wall 64 to form a tip cavity 70 with an open top end. The first tip wall 122 also extends a certain height (not shown) radially outward from the tip plate 54 to the outer edge 126. In the exemplary embodiment, the height of the first tip wall is equal to the height 74 of the second tip wall. Alternatively, the height of the first tip wall is not equal to the height 74 of the second tip wall.
[0025]
The first tip wall 122 is at least partially recessed from the airfoil first side wall 44. More specifically, the first tip wall 122 is recessed from the airfoil first side wall 44 toward the second tip wall 64, generally between the airfoil leading edge 48 and the trailing edge 50. A second tip shelf 130 that extends radially outward is formed. More specifically, the shelf 130 includes a front end edge 134 and a rear end edge 136. The front end edge 134 and the rear end edge 136 are each tapered and become the same plane as the first side wall 44. The shelf leading edge 134 is located downstream of the airfoil leading edge 48 by a distance 138, and the shelf trailing edge 136 is located upstream of the airfoil trailing edge 50 by a distance 140.
[0026]
The recessed first tip wall 122 and second tip shelf 130 form a substantially L-shaped trough 144 therebetween. In the exemplary embodiment, tip plate 54 is generally non-porous and includes a plurality of holes 106 extending through tip plate 54 in first tip shelf 90 and in second tip shelf 130. A plurality of holes 146 extending through the tip plate 54 are included. The holes 146 are axially spaced along the second tip shelf 130 and are in fluid communication between the trough 144 and the airfoil interior cooling chamber. In one embodiment, tip region 60 and airfoil 42 are coated with a thermal barrier coating.
[0027]
During operation, the squealer tip walls 122 and 64 are placed in close proximity to a conventional fixed stator shroud (not shown), forming a narrow gap (not shown) therebetween to prevent combustion gas leakage therethrough. Promote reduction. The tip wall 122 functions similarly to the tip wall 62 described above and extends radially outward from the airfoil 42. Thus, when rubbing occurs between the rotor blade 40 and the stator shroud, only the tip walls 122 and 64 are in contact with the shroud and the airfoil 42 is kept intact.
[0028]
In addition, the airfoil pressure side 46 and airfoil suction side 44 provided by the troughs 102 and 144, respectively, as the rotor blade 40 rotates and combustion gases flow past the airfoil tip shelves 90 and 130, respectively. Each discontinuity causes combustion gas to delaminate from the airfoil sidewalls 46 and 44, respectively, thus facilitating a reduction in its heat transfer. Trough 144 functions similarly to trough 102 and promotes film cooling circulation.
[0029]
FIG. 4 is a partial perspective view of another embodiment of a rotor blade 200 that may be used in a gas turbine engine, such as gas turbine engine 10 (shown in FIG. 1). The rotor blade 200 is substantially similar to the rotor blade 40 shown in FIG. 2, and components in the rotor blade 200 that are identical to the components of the rotor blade 40 are indicated using the same reference numerals used in FIG. 4 is specified. Accordingly, the rotor blade 200 includes an airfoil 42, sidewalls 44 and 46 extending between the leading edge 48 and the trailing edge 50, respectively, and a notch 80. Further, the rotor blade 200 includes a first tip wall 62, a notch 80, and a second tip wall 202. The notches 80 are formed between the first tip wall 62 and the second tip wall 202, respectively.
[0030]
The second tip wall 202 extends from the vicinity of the airfoil leading edge 48 to the airfoil trailing edge 50 along the airfoil second side wall 46. More specifically, the second tip wall 202 extends from the tip plate 54 to the outer edge 204 by a certain height (not shown). The height of the second tip wall is substantially constant along the second tip wall 202. The second tip wall 202 is spaced laterally from the first tip wall 62 to form a tip cavity 70 with an open top. In the exemplary embodiment, the height of the second tip wall is equal to the height 66 of the first tip wall. Alternatively, the height of the second tip wall is not equal to the height 66 of the first tip wall.
[0031]
The notch 80 includes a guide wall 210 that extends from the first tip wall 62 toward the trailing edge of the airfoil. More specifically, the guide wall 210 extends curvedly from the first tip wall 62 to form a curved inlet 212 for the notch 80. The guide wall 210 has a length 214 that is selected to allow the air flow entering the open top end cavity 70 toward the first end wall 62.
[0032]
The rotor blade described above is cost-effective and very reliable. The rotor blade includes a leading edge notch formed between the leading edges of the first and second tip walls. The tip walls are connected at the trailing edge of the rotor blade to form a tip cavity. In an exemplary embodiment, one of the tip walls is recessed to form a tip shelf. When the rotor blade rotates during operation, the tip wall prevents the rotor blade from rubbing against the stationary structural member. When combustion gases flow past the rotor blades, the rotor blade notch reduces heating of the tip cavity without increasing cooling air requirements or sacrificing rotor blade aerodynamic efficiency. Useful. In addition, the tip ledge disturbs the combustion gas flowing beyond the airfoil and promotes the formation of a cooling layer on the ledge. As a result, the lower operating temperature within the rotor blade can extend the useful life of the rotor blade in a cost-effective and reliable manner.
[0033]
While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.
[Brief description of the drawings]
FIG. 1 is a schematic view of a gas turbine engine.
FIG. 2 is a partial perspective view of a rotor blade that can be used in the gas turbine engine shown in FIG. 1;
FIG. 3 is a cross-sectional view of another embodiment of the rotor blade shown in FIG. 2;
4 is a partial perspective view of yet another embodiment of a rotor blade that can be used in the gas turbine engine shown in FIG. 1. FIG.
[Explanation of symbols]
42 Airfoil portion 44 First side wall 46 Second side wall 48 Leading edge 50 Trailing edge 54 Tip plate 60 Tip region 62 First tip wall 64 Second tip wall 70 Opening tip cavity 80 Notch 90 Tip shelf 102 Trough 106 Hole

Claims (9)

An airfoil (42) for a gas turbine engine (10) comprising:
The leading edge (48),
The trailing edge (50);
A tip plate (54);
A first sidewall (44) extending in a radial span between the airfoil root and the tip plate;
A second side wall (46) connected to the first side wall at the leading and trailing edges and extending in a radial span between the airfoil root and the tip plate;
A first tip wall (62) extending radially outward from the tip plate along the first side wall;
A second tip wall (64) extending radially outward from the tip plate along the second side wall and connected to the first tip wall at the trailing edge;
A notch (80) extending between the first tip wall and the second tip wall along the airfoil leading edge;
Only including, the notch (80),
An air flow entering the open top cavity (70) formed by the first and second tip walls (62, 64) and the tip plate (54) is directed toward the first tip wall (62); A guide wall (210) extending curvedly from the first tip wall toward the rear edge is formed, and the guide wall forms a curved entrance of the notch (80). Airfoil (42). Said second tip wall (64), the second is recessed from the sidewall (46) at least partially inwardly, characterized in that it forms a first tip shelf (90), wherein The airfoil (42) according to item 1 . The airfoil (42) of claim 1 , wherein the first tip wall (62) and the second tip wall (64) are substantially equal in height. The first tip wall (62) extends from the tip plate (54) by a first distance ( 66 ), and the second tip wall (64) extends from the tip plate by a second distance (74). The airfoil (42) according to claim 1, characterized in that: A gas turbine engine (10) comprising a plurality of rotor blades (40, 120, 200) comprising:
Each of the rotor blades includes a leading edge (48), a trailing edge (50), a first side wall (44), a second side wall (46), a first tip wall (62), a second tip wall (64), and Including an airfoil (42) including a notch (80);
The airfoil first and second sidewalls are connected axially at the leading and trailing edges and extend radially from a blade root to the tip plate (54);
The first tip wall extends radially outward from the tip plate along the first side wall;
The second tip wall extends radially outward from the tip plate along the second side wall;
The notch extends from the tip plate along the airfoil leading edge between the first tip wall and the second tip wall, and the notch (80) is
An air flow entering the open top cavity (70) formed by the first and second tip walls (62, 64) and the tip plate (54) is directed toward the first tip wall (62); A guide wall (210) extending curvedly from the first tip wall toward the rear edge is formed, and the guide wall forms a curved entrance of the notch (80). Gas turbine engine (10). The gas turbine engine of claim 5 , wherein the rotor blade airfoil first sidewall (44) is convex and the rotor blade airfoil second sidewall (46) is concave. (10). The second tip wall (64) is at least partially recessed inwardly from the second side wall (46) to form a first tip shelf (90). Item 7. The gas turbine engine (10) according to Item 5 or 6.   The gas turbine engine (10) according to any one of claims 5 to 7, wherein the first tip wall (62) and the second tip wall (64) are substantially equal in height. The first tip wall (62) extends from the tip plate (54) by a first distance ( 66 ), and the second tip wall (64) extends from the tip plate by a second distance (74). A gas turbine engine (10) according to any one of claims 5 to 8, characterized in that

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